A turbofan engine generally includes a fan assembly and a core gas turbine engine. The gas turbine engine includes, in serial flow order, a low pressure compressor, a high pressure compressor, a combustion section, a high pressure turbine and a low pressure turbine. A high pressure shaft couples the high pressure compressor to the high pressure turbine. A low pressure shaft extends coaxially within the high pressure shaft and couples the low pressure compressor to the low pressure turbine. The fan assembly includes a plurality of fan blades coupled to a fan shaft and disposed upstream from an inlet of the low pressure compressor. The fan shaft is coupled to the low pressure shaft via a gearbox. In particular configurations, an outer casing or nacelle circumscribes the fan blades and at least a portion of the gas turbine engine. A bypass air passage is defined between an outer casing of the gas turbine engine and the nacelle.
The combustion section generally includes an annular inner liner, an annular outer liner radially spaced from the inner liner and a combustor dome coupled to upstream or forward ends of the inner and outer liners. A fuel injector or nozzle extends through the dome and is configured to provide a fuel/air mixture to a combustion chamber that is defined between the inner and outer liners. An outer casing circumferentially surrounds the outer liner and at least partially defines an outer plenum or passage therebetween.
The combustion section further includes an ignition system having one or more igniter assemblies mounted or coupled to the outer casing. An igniter portion of the igniter assembly extends generally radially through the outer casing and the outer plenum. An ignition tip portion of the igniter extends at least partially through an opening defined within the outer liner. During operation of the gas turbine, such as during light-off or restart, the igniter may be energized to provide a spark at the ignition tip so as to ignite the fuel/air mixture within the combustion chamber.
Radial and/or axial positioning of the ignition tip with respect to the outer liner and/or the combustion chamber may change during operation of the gas turbine. For example, varying thermal growth rates of the outer casing and the outer liner and/or g-forces may result in over immersion of the ignition tip into the flow of extremely hot combustion gases, thus resulting in undesirable thermal fatigue of the igniter. In addition or in the alternative, varying thermal growth rates of the outer casing and the outer liner and/or g-forces may cause the ignition tip to lift radially out of the opening defined within the outer liner, thus potentially affecting the ability to light off the combustor. Consequently, an improved ignition system for a gas turbine engine would be useful in the turbofan engine industry.